The disclosure relates generally to turbine systems, and more particularly, to reducing pressure loss in a multi-wall turbine blade cooling circuit.
Gas turbine systems are one example of turbomachines widely utilized in fields such as power generation. A conventional gas turbine system includes a compressor section, a combustor section, and a turbine section. During operation of the gas turbine system, various components in the system, such as turbine blades, are subjected to high temperature flows, which can cause the components to fail. Since higher temperature flows generally result in increased performance, efficiency, and power output of the gas turbine system, it is advantageous to cool the components that are subjected to high temperature flows to allow the gas turbine system to operate at increased temperatures.
Turbine blades of a gas turbine system typically contain an intricate maze of internal cooling channels. The cooling channels receive air from the compressor of the gas turbine system and pass the air through the internal cooling channels to cool the turbine blades. The teed pressure of the air passed through the cooling channels is generally at a premium, since the air is bled off of the compressor. To this extent, it is useful to provide cooling channels that reduce non-recoverable pressure loss; as pressure losses increase, a higher feed pressure is required to maintain an adequate gas-path pressure margin (back-flow margin). Higher feed pressures result in higher leakages in the secondary flow circuits (e.g., in rotors) and higher feed temperatures.